Cooling system for a turbine engine

ABSTRACT

A gas turbine engine including a compressor section, a turbine section, and a combustion section positioned between the compressor section and the turbine section is provided. The gas turbine engine also includes a cooling system having a tank and one or more fluid lines in fluid communication with the tank. The one or more fluid lines are configured to carry a flow of consumable cooling liquid provide such consumable cooling liquid to one or more components of the compressor section, the turbine section, and/or the combustion section not directly exposed to a core air flowpath defined through the gas turbine engine.

FIELD OF THE INVENTION

The present subject matter relates generally to a cooling system for agas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine general includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gassesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

It is desirable to operate the gas turbine engine at relatively hightemperatures and pressures to increase an efficiency of the gas turbineengine or to maximize a power output of the gas turbine engine. Forexample, it is especially desirable to operate the gas turbine engine atrelatively high temperatures and pressures during modes of operationrequiring high power output, such as during takeoff and climb of anaircraft having such an engine. However, in order to withstand suchrelatively high temperatures and pressures, various components of thegas turbine engine must be constructed of rare and/or expensivematerials.

Accordingly, a device for cooling various components of the gas turbineengine without substantially reducing an operating temperature and/orpressure of the gas turbine engine would be useful. More particularly, adevice for cooling various components of the gas turbine engine notdirectly exposed to the core air flowpath of the gas turbine enginewould be particularly beneficial.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a gas turbineengine defining an axial direction is provided. The gas turbine engineincludes a compressor section, a turbine section located downstream ofthe compressor section, and a combustion section positioned between thecompressor section and the turbine section. The compressor section, theturbine section, and the combustion section together define a core airflowpath. The gas turbine engine also includes a cooling system forcooling one or more components of the compressor section, the turbinesection, or the combustion section not directly exposed to the core airflowpath. The cooling system includes a fluid tank for storing a volumeof consumable cooling liquid, and one or more fluid lines. The one ormore fluid lines are in fluid communication with the fluid tank forcarrying a flow of the consumable cooling liquid and providing theconsumable cooling liquid to the one or more components of thecompressor section, the turbine section, or the combustion section notdirectly exposed to the core air flowpath.

In one exemplary aspect of the present disclosure, a method for coolinga gas turbine engine defining a core air flowpath extending through acompressor section, a combustion section, and a turbine section isprovided. The method includes operating the gas turbine engine toachieve a high power output, and providing a consumable cooling liquidusing a cooling system to a surface of one or more components of thecompressor section, the combustion section, or the turbine section, thesurface not directly exposed to the core air flowpath.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a schematic cross-sectional view of an exemplary gas turbineengine according to various embodiments of the present subject matter.

FIG. 2 is close-up, simplified, schematic view of the exemplary gasturbine engine of FIG. 1 including a cooling system in accordance withan exemplary embodiment of the present disclosure.

FIG. 3 is a close-up, simplified, schematic view of an aft end of acompressor section of the exemplary gas turbine engine of FIG. 1 and aportion of the exemplary cooling system depicted in FIG. 2.

FIG. 4 is a schematic view of the exemplary cooling system depicted inFIG. 2, taken along an axial direction of the exemplary gas turbineengine of FIG. 1.

FIG. 5 is a close-up, simplified, schematic view of the exemplary gasturbine engine of FIG. 1 including a cooling system in accordance withanother exemplary embodiment of the present disclosure.

FIG. 6 is a close-up, simplified, schematic view of an aft end of acompressor section of the exemplary gas turbine engine of FIG. 1 and aportion of the exemplary cooling system depicted in FIG. 5.

FIG. 7 is a close-up, simplified, schematic view of the exemplary gasturbine engine of FIG. 1 including a cooling system in accordance withyet another exemplary embodiment of the present disclosure.

FIG. 8 is a close-up, simplified, schematic view of the exemplary gasturbine engine of FIG. 1 including a cooling system in accordance withstill another exemplary embodiment of the present disclosure.

FIG. 9 is a flow diagram of an exemplary method for operating a gasturbine engine.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first”, “second”, and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows, and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference), a radial direction R, and a circumferential direction C (seeFIG. 4). In general, the turbofan 10 includes a fan section 14 and acore turbine engine 16 disposed downstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22. The compressor section,combustion section 26, and turbine section together define a core airflowpath 37.

For the embodiment depicted, the fan section 14 includes a variablepitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 ina spaced apart manner. As depicted, the fan blades 40 extend outwardlyfrom the disk 42 generally along the radial direction R. Each fan blade40 is rotatable relative to the disk 42 about a pitch axis P by virtueof the fan blades 40 being operatively coupled to a suitable actuationmember 44 configured to collectively vary the pitch of the fan blades 40in unison. The fan blades 40, disk 42, and actuation member 44 aretogether rotatable about the longitudinal axis 12 by LP shaft 36 acrossa power gear box 46. The power gear box 46 includes a plurality of gearsfor stepping down the rotational speed of the LP shaft 36 to a moreefficient rotational fan speed.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 iscovered by rotatable front hub 48 aerodynamically contoured to promotean airflow through the plurality of fan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing or outer nacelle50 that circumferentially surrounds the fan 38 and/or at least a portionof the core turbine engine 16. It should be appreciated that the nacelle50 may be configured to be supported relative to the core turbine engine16 by a plurality of circumferentially-spaced outlet guide vanes 52.Moreover, a downstream section 54 of the nacelle 50 may extend over anouter portion of the core turbine engine 16 so as to define a bypassairflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58, entersthe turbofan 10 through an associated inlet 60 of the nacelle 50 and/orfan section 14. As the volume of air 58 passes across the fan blades 40,a first portion of the air 58, as indicated by arrows 62, is directed orrouted into the bypass airflow passage 56 and a second portion of theair 58, as indicated by arrow 64, is directed or routed into the coreair flowpath 37, or more particularly, into the LP compressor 22. Theratio between the first portion of air 62 and the second portion of air64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

It should be appreciated, however, that the exemplary turbofan engine 10depicted in FIG. 1 is by way of example only, and that in otherexemplary embodiments, aspects of the present disclosure mayadditionally, or alternatively, be applied to any other suitable gasturbine engine. For example, in other exemplary embodiments, the gasturbine engine may be any other suitable aeronautical gas turbineengine, such as a turbojet engine, turboshaft engine, turboprop engine,etc.

Referring now to FIG. 2, a close-up, side, schematic view of theturbofan engine 10 of FIG. 1 is provided. More particularly, FIG. 2provides a close-up view of an aft end of the compressor section (ormore particularly, an aft end 80 of the HP compressor 24), thecombustion section 26, and a forward end 82 of the turbine section (ormore particularly, a forward end 82 of the HP turbine 28). Additionally,as will be discussed in greater detail below, the turbofan engine 10includes a cooling system 100 in accordance with an exemplary embodimentof the present disclosure for cooling compressed air in or from thecompressor section and/or for cooling certain components of thecompressor section proximate to the aft end 80 of the HP compressor 24.

As depicted, the HP compressor 24 includes various stages of rotorblades 102 rotatable about the longitudinal centerline 12. Each rotorblade 102 in a particular stage is attached at a root 104 to one of aplurality HP compressor rotors 106. Each HP compressor rotor 106includes a bore 108 and a web 110 and is connected to an adjacent HPcompressor rotor 106 via a spacer arm 112. Each stage of rotor blades102 progressively compresses the air flowing through that portion of thecore air flowpath 37. Disposed between each adjacent stage of rotorblades 102 in the HP compressor 24 is a stage of stationary stator vanes114 attached to an HP compressor liner 116. An inner surface of the HPcompressor liner 116 is exposed directly to and at least partiallydefines the core air flowpath 37. Each stator vane 114 includes a seal118 at a radially inner end 120, positioned adjacent to a plurality ofteeth 122 extending generally along the radial direction R from thespacer arms 112. The seal 118 and teeth 122 are configured to reduce anamount of airflow allowable around the radially inner ends 118 of thestator vane 114.

After the compressed air flows past an aft-stage 124 of rotor blades102, the compressed air flows through a diffuser 126 for channeling thecompressed air and directing such compressed air into the combustionsection 26. For the embodiment depicted, the diffuser 126 is furtherconfigured to reduce a Mach number of the flow of compressed air suchthat the fuel may more easily be ignited in the combustion section 26.

The compressed air from the HP compressor 24 is mixed with fuel usingone of a plurality of fuel-air mixers 128 of the combustion section 26.The combustion section 26 additionally includes an inner liner 130 andan outer liner 132, the inner and outer liners 130, 132 together atleast partially defining a combustion chamber 134. The mixture of fueland compressed air is combusted in the combustion chamber 134 togenerate the combustion gases, which flow from the combustion chamber134 into the HP turbine 28. The flow of combustion gases through the HPturbine 28 rotates the sequential stages of HP turbine rotor blades 70.Disposed between or adjacent to each stage of HP turbine rotor blades 70is stage of turbine stator vanes 68 attached to a liner 135. An innersurface of the HP turbine liner 135 is exposed directly to and at leastpartially defines the core air flowpath 37. Additionally, each HPturbine rotor blades 70 in a particular stage of HP turbine rotor blades70 is attached at a base 138 to a respective HP turbine rotor 136.Rotation of the HP turbine blades 70 and a respective HP turbine rotor136 drives the HP turbine 28. The HP turbine rotor 136 is coupled withthe HP rotor shaft 34, which is also coupled to the HP compressor rotor106 through an arm 140 of the HP rotor shaft 34. Accordingly, asdiscussed, rotation of the HP turbine rotor blades 70 drives the HPcompressor 24.

Notably, the HP rotor shaft 34, for the embodiment depicted,additionally includes a compressor discharge pressure seal 142, whichfor the embodiment depicted is configured for metering a flow ofcompressed air from the HP compressor 24 to a radially inner cavity 144defined between the HP rotor shaft 34 and a forward inner nozzle supportmember 146.

As stated, the exemplary turbofan engine 10 depicted in FIG. 2additionally includes the exemplary cooling system 100. The exemplarycooling system 100 generally includes a fluid tank 148 for storing avolume of cooling fluid and one or more fluid lines. The one or morefluid lines are in fluid communication with the fluid tank 148 forinjecting cooling fluid into, e.g., the compressed air proximate theaft-stage 124 of rotor blades 102 of the HP compressor 24. Morespecifically, for the exemplary cooling system 100 depicted, the one ormore fluid lines include a supply fluid line 150 in fluid communicationwith the tank 148, a compressor fluid line 152, and a turbine fluid line154. Additionally, the exemplary cooling system 100 includes a valve156, the valve 156 including an inlet fluidly connected with the supplyfluid line 150, a first outlet fluidly connected to the compressor fluidline 152, and a second outlet fluidly connected to the turbine fluidline 154.

The valve 156 may be a standard three-way valve providing a fixed ratioof cooling fluid from the inlet to the first outlet and second outlet(i.e., from the supply fluid line 150 to the compressor fluid line 152and the turbine fluid line 154). Alternatively, however, the valve 156may be a variable throughput, three-way valve configured to alter theratio of cooling fluid provided from the inlet to the first outlet andsecond outlet. For example, the valve 156 may be configured to provideanywhere from zero percent (0%) to one hundred percent (100%) of thecooling fluid from the inlet to the first outlet. Additionally, thevalve 156 may be configured to provide anywhere from zero percent (0%)to one hundred percent (100%) of the cooling fluid from the inlet to thesecond outlet. Moreover, the valve 156 may be capable of restricting atotal amount of cooling fluid allowable therethrough to both the firstand second outlets. Accordingly, the valve 156 may be capable ofeffectively shutting off a flow of cooling fluid to the compressor fluidline 152 and/or the turbine fluid line 154 based on, e.g., a need forsuch cooling.

In certain exemplary embodiments, the cooling fluid may be a consumablecooling liquid, such as water or a glycol-water mixture (which may beresistant to, e.g., freezing). Alternatively, however, in otherembodiments, any other suitable cooling fluid may be provided. Notably,as used herein, the term “consumable” with reference to the coolingfluid refers to the manner in which the cooling fluid reduces atemperature of a component. Specifically, in at least certain exemplaryaspects, the term consumable refers to a liquid which is configured tovaporize after contacting a component, absorbing heat and reducing atemperature of such component.

As is also depicted in FIG. 2, the exemplary cooling system 100 furtherincludes a pump 158 for generating a flow of cooling fluid from the tank148 through a fluid line, such as the supply fluid line 150. The pump158 may be a rotary pump including an impeller, or alternatively, may beany other suitable fluid pump. Additionally, for the embodimentdepicted, the pump 158 and the fluid tank 148 are positioned outward ofthe core air flowpath 37 along the radial direction R. Accordingly, atleast one of the one or more fluid lines extends through the core airflowpath 37 to a location inward of the core air flowpath 37 along theradial direction R. For the embodiment depicted, a separate tube 160extending through the core air flowpath 37 is provided for housing thefluid line(s) of the cooling system 100 extending through the core airflowpath 37. Notably, in certain exemplary embodiments, the pump 158and/or the valve 156 may be in operable communication with a controllerof the turbofan engine 10 to control operation of the cooling system100. For example, the controller may regulate a flow rate of the coolingfluid through the cooling system 100 based on, e.g., an operatingcondition of the turbofan engine 10, or in response to one or moretemperature sensors (not shown).

It should be appreciated, however, that in other exemplary embodiments,the pump 158 and/or tank 148 may alternatively be positioned inward fromthe core air flowpath 37 along the radial direction R. In such aconfiguration, or in other configurations, the pump 158 may be directlyand/or indirectly coupled to, e.g., the HP shaft 34 and driven by, e.g.,the HP shaft 34.

Referring still to FIG. 2, the turbine fluid line 154 is configured forinjecting cooling fluid into a cooling airflow provided to the turbinesection. Specifically, for the embodiment depicted, the turbine fluidline 154 is configured for injecting cooling fluid into a coolingairflow flowing from the HP turbine 28 and through a cooling channel162. For the embodiment depicted, the cooling channel 162 is at leastpartially defined by the forward inner nozzle support 146. Additionally,the cooling channel 162 includes an axial inducer 164, which brings thecooling airflow from a static frame of reference to a rotating frame ofreference while minimizing a temperature rise due to stagnation onto theHP turbine rotor 136. The cooling airflow provided through the coolingchannel 162 may be used as cooling air provided to reduce a temperatureof certain components in the HP turbine 28. For example, the coolingairflow provided through the cooling channel 162 may be provided to theHP turbine rotor blades 70, such as a first stage 166 of HP turbinerotor blades 70, and/or as cavity purge air for a cavity of the HPturbine rotor 136.

As shown, the turbine fluid line 154 extends from the valve 156 andthrough at least a portion of the cooling channel 162 towards the HPturbine 28. The turbine fluid line 154 includes a nozzle 168 defining anoutlet for injecting cooling fluid into the cooling air flowing throughthe cooling channel 162. The cooling fluid injected into the cooling airflowing through the channel 162 may reduce a temperature of the coolingair, such that less cooling air is required to maintain a desiredtemperature of certain components in the HP turbine 28 and/or such thata temperature of certain components in the HP turbine 28 may be furtherreduced to increase a life of such components. Notably, if less coolingair is required to maintain a desired temperature of certain componentsin the HP turbine 28, less air is required to be taken from thecompressor section, such that an efficiency of the turbofan engine 10may be increased. It should be appreciated, however, that in otherexemplary embodiments, the turbine fluid line 154 may extend to anyother suitable location to decrease a temperature of cooling airprovided to, e.g., the HP turbine 28. For example, in other exemplaryembodiments, the turbine fluid line 154 may extend through the forwardinner nozzle support 146 for indirectly spraying cooling fluid onto theone or more HP turbine rotors 136.

It should also be appreciated that such a configuration differs fromprior art configurations that have included fluid injection systems inthe core air flowpath 37 extending through the turbine section. Suchprior art configurations are set up to reduce an overall temperature ofthe airflow in the core air flowpath 37 extending through the turbinesection. By contrast to these prior art configurations, the presentconfiguration may inject cooling fluid into the cooling airflow providedto certain components of the turbine section for reducing a temperatureof the cooling airflow itself. As will be discussed below, theconfiguration disclosed herein requires less cooling fluid than theprior art configurations.

Referring now also to FIG. 3, a close-up view of the aft end 80 of theHP compressor 24 is provided. Notably, one measure of an efficiency ofthe turbofan engine 10 is an overall pressure ratio of the turbofanengine 10. The overall pressure ratio refers generally to a ratio of apressure at the aft end 80 of the compressor section to a stagnationpressure at the forward end of the compressor section. However, as theoverall pressure ratio of the turbofan engine 10 increases, atemperature of the compressed air and components at the aft end 80 ofthe HP compressor 24 also increases. In order to allow for an increasedoverall pressure ratio without damaging the HP compressor 24, theexemplary cooling system 100 depicted in FIGS. 2 and 3 includes thecompressor fluid line 152 for cooling one or both of the compressed airin or from the aft end 80 of the HP compressor 24 or certain componentsof the HP compressor 24 at the aft end 80 of the HP compressor 24.

Specifically, the compressor fluid line 152 is configured for injectingcooling fluid into the compressed air proximate the aft-stage 124 ofrotor blades 102 of the HP compressor 24. The compressor fluid line 152includes an outlet, or rather includes a nozzle 170 defining an outlet,positioned adjacent to the aft-stage 124 of rotor blades of the HPcompressor 24. As used herein, “positioned adjacent to the aft-stage 124of rotor blades of the HP compressor 24” refers to being positionedclose enough to inject a flow of cooling fluid into the compressed airin the aft end 80 of the HP compressor 24. For the exemplary embodimentdepicted, the compressor fluid line 152 is attached to a stationaryframe member, inward of the core air flowpath 37 along the radialdirection R, at a location immediately downstream of the aft-stage 124of rotor blades 102. More particularly, for the embodiment depicted, thecompressor fluid line 152 is attached to a forward end 172 of theforward inner nozzle support 146, at a location inward of the diffuser126 along the radial direction R. With such a configuration, the outletdefined by the nozzle 170 is located immediately downstream of theaft-stage 124 of rotor blades 102. However, in other exemplaryembodiments, the compressor fluid line 152 may alternatively be attachedto any other suitable stationary member allowing the nozzle to inject aflow of cooling fluid to the aft end 80 of the HP compressor 24.

As stated, for the embodiment depicted, the nozzle 168 is positionedimmediately downstream of the aft-stage 124 of rotor blades 102.Accordingly, to inject a flow of cooling fluid to the aft end 80 of theHP compressor 24, the nozzle 168 is configured to spray the coolingfluid through the outlet generally towards a forward end of the HPcompressor 24 (i.e., in an upstream direction). Notably, duringoperation, i.e., when the aft-stage 124 of rotor blades 102 are rotatingabout the axial direction A, the high pressure, compressed air may tryto flow to areas of lower pressure. For example, the high pressure,compressed air may try to flow around the radially inner end 120 of thestator vanes 114 to an upstream location. Accordingly, the airflowthrough the aft end 80 of the HP compressor 24 tends to swirl around inthe area adjacent to the seal 118 of the stator vane 114 and theplurality of teeth 122 of the HP compressor rotor 106. As the overallpressure ratio of the turbofan engine 10 is increased, a temperature ofthese components may also increase—more so than other components—due tothe swirling of such high temperature compressed air. Accordingly, byinjecting the cooling fluid into the compressed air proximate theaft-stage 124 of rotor blades 102, the cooling fluid can travel forwardto the areas of increased temperatures (i.e., where the compressed airis swirling) and vaporize. The vaporization of the cooling fluid absorbsheat and cools the components and/or the compressed air.

Thus, a gas turbine engine having a cooling system 100 in accordancewith an exemplary embodiment present disclosure may be capable ofincreasing its overall pressure ratio without risk of damaging certaincomponents at an aft end of a compressor section. Accordingly, a gasturbine engine having a cooling system 100 in accordance with anexemplary embodiment of the present disclosure may be capable ofachieving a higher efficiency, while increasing a lifespan of certaincomponents at an aft end of the compressor section.

Referring now also to FIG. 4, a simplified schematic view along theaxial direction A of the turbofan engine 10 is provided of the exemplarycooling system 100 of FIG. 2. As shown, for the embodiment depicted, theone or more fluid lines further includes a plurality of compressor fluidlines 152. The plurality of compressor fluid lines 152 each include anoutlet, or rather a nozzle 170 defining an outlet, with the nozzles 170and outlets of the plurality of fluid lines 152 spaced around theturbofan engine 10 about the circumferential direction C of the turbofanengine 10. The outlet of each of the plurality of compressor fluid lines152 may be positioned adjacent to the aft-stage 124 of rotor blades 102of the HP compressor 24 for injecting cooling fluid into the compressedair proximate the aft-stage 124 rotor blades 102. For the embodimentdepicted, the plurality of compressor fluid lines 152 includes sixteen(16) compressor fluid lines 152 circumferentially spaced about the axialdirection A and positioned adjacent to the aft-stage 124 rotor blades102 of the HP compressor 24. However, in other exemplary embodiments,the plurality of compressor fluid lines 152 may include less thansixteen (16) compressor fluid lines 152, or alternatively the pluralityof fluid lines may include at least twenty (20) compressor fluid lines152, at least twenty-five (25) compressor fluid lines 152, or at leastthirty (30) compressor fluid lines 152.

It should be appreciated, that in certain exemplary embodiments, thecooling system 100 may additionally include a similar configuration ofturbine fluid lines 154. More particularly, in certain exemplaryembodiments, the cooling system 100 may additionally include a pluralityof turbine fluid lines 154 spaced along the circumferential direction Cin, e.g., the cooling airflow channel 162. However, in other exemplaryembodiments, the cooling system 100 may have any other suitableconfiguration for the turbine fluid lines 154. Alternatively, in stillother exemplary embodiments, the cooling system 100 may not include anyturbine fluid lines 154, and instead may be focused on providing acooling fluid to, e.g., the aft end 80 of the HP compressor 24.

In certain embodiments, the cooling system 100 may be configured toinject up to two pounds of cooling fluid per second into the compressedair proximate the aft-stage 124 of rotor blades 102 of the HP compressor24 (e.g., through the compressor fluid line 152). Alternatively,however, the cooling system 100 may be configured to inject up to aboutthree pounds of cooling fluid per second, up to about four pounds ofcooling fluid per second, up to about five pounds of cooling fluid persecond, or up to about six pounds of cooling fluid per second into thecompressed air proximate the aft-stage 124 of rotor blades 102 of the HPcompressor 24. It should be appreciated, that as used herein, terms ofapproximation, such as “about” or “approximately,” refer to being withina ten percent (10%) margin of error.

Additionally, or alternatively, the cooling system 100 may be configuredto inject cooling fluid into the compressed air proximate the aft-stage124 of rotor blades 102 of the HP compressor 24 (e.g., through thecompressor fluid line 152) at a rate greater than about 0.05% of a massflow rate of the compressed air flowing through the HP compressor 24 andless than about ten percent (10%) the mass flow rate of the compressedair flowing through the HP compressor 24. However, in other embodiments,the cooling system 100 may be configured to inject between about 0.05%and about five percent (5%), between about 0.05% and about three percent(3%), or between about 0.05% and about two percent (2%) of the mass flowrate of the compressed air flowing through the HP compressor 24.

Notably, the cooling system 100 may additionally, or alternatively, beconfigured to inject cooling fluid into, e.g., the cooling airflowthrough the turbine fluid line(s) 154 at a same or similar rate asthrough the compressor fluid line(s) 152 and/or at a same or similarratio as through the compressor fluid line(s) 152.

Further, in certain exemplary aspects, the cooling system 100 may onlybe operated during times wherein an increased overall pressure ratio isdesired. For example, as discussed below with respect to FIG. 9, thecooling system 100 may only be operated to cool the aft end 80 of the HPcompressor 24 during periods of peak power of the turbofan engine 10,such as when an aircraft having such an exemplary turbofan engine 10 istaking off or climbing.

As stated, a cooling system 100 in accordance certain exemplaryembodiments of the present disclosure may be capable of cooling at leastone of the airflow proximate the aft-stage 124 of rotor blades in the HPcompressor 24, or certain components of the HP compressor 24 proximatethe aft-stage 124 of rotor blades in the HP compressor 24. Thus, acooling system 100 in accordance with certain exemplary embodiments ofthe present disclosure may allow for an increased overall pressure ratioof the gas turbine engine.

Referring now to FIGS. 5 and 6, an alternative embodiment of theexemplary cooling system 100 described above with reference to FIGS. 2through 4 is provided. More particularly, FIG. 5 provides a simplifiedschematic view of the exemplary turbofan engine 10 of FIG. 1, includinga cooling system 100 in accordance with another exemplary embodiment ofthe present disclosure; and FIG. 6 provides a close-up, schematic viewof the aft end 80 of the HP compressor 24 of the turbofan engine 10 ofFIG. 1 including the exemplary cooling system 100 of FIG. 5.

As stated, in other exemplary embodiments of the cooling system 100, thecompressor fluid line 152 may be positioned at any suitable location toallow the nozzle 170 of the compressor fluid line 152 to inject a flowof cooling fluid to the aft end 80 of the HP compressor 24. Theexemplary cooling system 100 of FIGS. 5 and 6 may be configured insubstantially the same manner as the exemplary cooling system 100described above with reference to FIGS. 2 through 4. However, for theembodiment depicted in FIGS. 5 and 6, the compressor fluid line 152 ismounted adjacent to or integrated into a compressor stator vanes 114 atthe aft end 80 of the HP compressor 24. More particularly, for theembodiment depicted, the compressor fluid line 152 is fluidly connectedto a fluid conduit extending through the stator vane 114. For thepurposes of this disclosure, the fluid conduit extending through thestator vane 114 is considered the nozzle 170 of the compressor fluidline 152. The fluid conduit in the stator vane 114 defines an outletproximate to the radially inner end 120 for injecting a flow of coolingfluid into the airflow at the aft end 80 of the HP compressor 24.Notably, with such a configuration, the three-way valve 156 is locatedoutward of the core air flowpath 37 along the radial direction R.

It should be appreciated, however, that in other exemplary embodiments,the compressor fluid line 152 may additionally, or alternatively, bepositioned at any other suitable location for injecting cooling fluid tothe aft end of the HP compressor 24. For example, in other exemplaryembodiments, a portion of the compressor fluid line 152 may be attachedto, and extend along, a surface of the stator vane 114, with the nozzle170 positioned proximate the radially inner end 120 of the stator vane114. As with the exemplary embodiment of FIGS. 2 through 4, theexemplary cooling system 100 wall FIGS. 5 and 6 may include a pluralityof compressor fluid lines 152 integrated into, or positioned adjacentto, stator vanes 114 and spaced along the circumferential direction C ofthe turbofan engine 10.

Reference will now be made to further exemplary embodiments of thepresent disclosure. For example, it should be appreciated that in stillother exemplary embodiments of the present disclosure, the coolingsystem 100 may be configured to cool any other suitable components ofthe turbofan engine 10. For example, in other exemplary embodiments,such as with the various exemplary embodiments discussed below withreference to FIGS. 7 and 8, the cooling system 100 may additionally, oralternatively, be configured for cooling the one or more components ofthe compressor section, the turbine section, or the combustion section26 not directly exposed to the core air flowpath 37. With such aconfiguration, the one or more fluid lines maybe configured for carryinga flow of the cooling fluid and providing the cooling fluid directly orindirectly to the one or more components of the compressor section,turbine section, or combustion section 26 not directly exposed to thecore air flowpath 37.

Reference will now be made to FIG. 7, providing a simplified, schematicview of the turbofan engine 10 of FIG. 1 including a cooling system 100in accordance with yet another exemplary embodiment of the presentdisclosure. The exemplary cooling system 100 of FIG. 7 may be configuredin substantially the same manner as the exemplary cooling system 100described above with reference to FIGS. 2 through 4. However, for theembodiment of FIG. 7, the one or more fluid lines include a compressorportion 174 for providing the cooling fluid to a liner of the compressorsection, a combustion portion 176 for providing the cooling fluid to aliner of the combustion section 26, and a turbine portion 178 forproviding the cooling fluid to a liner of the turbine section. Moreparticularly, for the embodiment of FIG. 7, the compressor portion 174of the one or more fluid lines is positioned adjacent to an outersurface 180 of the compressor liner 116 for spraying the cooling fluidon the outer surface 180 of the compressor liner 116. Additionally, thecombustion portion 176 of the one or more fluid lines is positionedadjacent to an outer surface 182 of the outer combustion chamber liner132 for spraying the cooling fluid on the outer surface 182 of the outercombustion chamber liner 132. Moreover, the turbine portion 178 of theone or more fluid lines is positioned adjacent to an outer surface 184of the turbine liner 135 for spraying the cooling fluid on the outersurface 184 of the turbine liner 135.

For the embodiment depicted, the compressor portion 174 of the one ormore fluid lines includes a plurality of nozzles 186, each nozzle 186defining an outlet for spraying the cooling fluid onto the outer surface180 of the compressor liner 116. Similarly, the combustion portion 176of the one or more fluid lines also includes a plurality of nozzles 188,each nozzle 188 defining an outlet for spraying the cooling fluid ontothe outer surface 182 of the outer combustion chamber liner 132.Further, the turbine portion 178 of the one or more fluid linessimilarly includes a plurality of nozzles 190, each nozzle 190 definingan outlet for spraying the cooling fluid onto the outer surface 184 ofthe turbine liner 135. The plurality of nozzles 186, 188, 190 in thecompressor portion 174, the combustion portion 176, and the turbineportion 178, respectively, of the one or more fluid lines are allgenerally spaced along the axial direction A.

It should be appreciated, however, that in other exemplary embodiments,the compressor portion 174, the combustor portion 176, and/or theturbine portion 178 of the one or more fluid lines may additionally, oralternatively, be configured for providing the cooling liquid indirectlyto the respective compressor liner 116, outer combustion chamber liner132, and turbine liner 135. For example, in other exemplary embodiments,the compressor portion 174, the combustor portion 176, and/or theturbine portion 178 of the one or more fluid lines may be configured forindirectly providing the cooling liquid to such components byspraying/injecting the cooling liquid into a flow of cooling airprovided over such components. Such a flow of cooling air may beextracted from the compressor section of the turbofan engine 10.

Referring still to the embodiment depicted in FIG. 7, the compressorportion 174 of the one or more fluid lines is fluidly connected to thesupply fluid line 150 via a valve 156. Similarly, the combustion portion176 of the one or more fluid lines is fluidly connected to the supplyfluid line 150 via the valve 156. Additionally, the turbine portion 178of the one or more fluid lines is fluidly connected to the combustionportion 176 of the one or more fluid lines. In other exemplaryembodiments, however, each of the compressor portion 174, combustionportion 176, and turbine portion 178 may be directly fluidly connectedto the supply fluid line 150 via valve 156, and thus the exemplary valve156 may be a four-way valve. In such an embodiment, the valve 156 may beconfigured to independently control an amount of cooling liquid providedthrough each of the compressor portion 174, combustion portion 176, andturbine portion 178 of the one or more fluid lines. Alternatively,however, each of the compressor portion 174, combustion portion 176, andturbine portion 178 may be configured in series flow with one another.

Additionally, although not depicted, in certain exemplary embodimentsthe one or more fluid lines, including each of the compressor portion174, combustion portion 176, and turbine portion 178, may include aplurality of fluid lines with at least a portion of each of theplurality of fluid lines spaced generally along the circumferentialdirection C of the turbofan engine 10 within a casing 192 of theturbofan engine 10. For example, the plurality of fluid lines may eachbe spaced along the circumferential direction C in substantially thesame manner that the plurality of compressor fluid lines 152 are spacedalong the circumferential direction C, as described above with respectto FIG. 4. Such a configuration may provide for an even cooling fluiddistribution over the outer surfaces 180, 182, 184 of the compressorliner 116, outer combustion chamber liner 132, and/or turbine liner 135,respectively, generally along the circumferential direction C. Moreparticularly, such a configuration may provide for a substantially eventemperature reduction across the compressor liner 116, the outercombustion chamber liner 132, and/or the turbine liner 135 generallyalong the circumferential direction C.

Notably, the exemplary cooling system 100 of FIG. 7 may be configured toinject cooling fluid through each of the compressor portion 174, thecombustion portion 176, and/or the turbine portion 178, individually orcumulatively, at the same rate and/or the same ratio that the exemplarycooling system 100 described above with reference to FIGS. 2 through 4is configured to provide cooling fluid through the compressor fluidline(s) 152.

Referring now to FIG. 8, a cooling system 100 in accordance with yetanother exemplary embodiment of the present disclosure is provided. FIG.8 provides a simplified, schematic view of the turbofan engine 10 ofFIG. 1 including a cooling system 100 in accordance with still anotherexemplary embodiment of the present disclosure.

The exemplary cooling system 100 of FIG. 8 is configured insubstantially the same manner as the exemplary cooling system 100 ofFIG. 7. Specifically, the exemplary cooling system 100 of FIG. 8 isconfigured for cooling one or more components of the compressor section,the turbine section, or the combustion section 26 not directly exposedto the core air flowpath 37. The exemplary cooling system 100 of FIG. 8,however, is alternatively configured to cool one or more of suchcomponents located inward of the core air flowpath 37 along the radialdirection R.

Specifically, for the embodiment of FIG. 8, the cooling system 100includes one or more fluid lines configured for providing the coolingfluid to a surface of the plurality of rotors 106 of the compressorsection, a surface of the plurality of rotors 136 of the turbinesection, and a surface of the compressor discharge pressure seal 142.Specifically, for the embodiment of FIG. 8, the one or more fluid linesinclude at least a first, outer fluid line 194 and a second, inner fluidline 196. The outer fluid line 194 is connected to the supply fluid line150 via valve 156. Additionally, the exemplary cooling system 100includes a static to rotating frame fluid transfer mechanism 198attached to the arm 140 of the HP rotor shaft 34 and fluidly connectedto the one or more fluid lines. Specifically, for the embodimentdepicted, the outer fluid line 194 is fluidly connected to the innerfluid line 196 through the static to rotating frame fluid transfermechanism 198. The inner fluid line 196 includes a plurality ofnozzles—a first nozzle 200, a second nozzle 202, and a third nozzle204—each nozzle defining an outlet. An outlet of the first nozzle 200 ispositioned proximate to and is directed towards a surface of at leastone of the plurality of HP compressor rotors 106. An outlet of thesecond nozzle 202 is positioned proximate to and is directed towards asurface of the compressor discharge pressure seal 142. Additionally, anoutlet of the third nozzle 204 is positioned proximate to and isdirected towards a surface of at least one of the plurality of HPturbine rotors 136.

In certain embodiments, the static to rotating frame fluid transfermechanism 198 may be configured as one or more journal bearings operablewith a plurality of circumferentially spaced and radially extendingholes in the HP rotor shaft 34. For example, the static to rotatingframe fluid transfer mechanism 198 may include an outer journal bearingextending around an outside surface 206 of the arm 140 of the HP rotorshaft 34. Specifically, the outer journal bearing may be positioned overa portion of the arm 140 of the HP rotor shaft 34 including theplurality of circumferentially spaced, radially extending holes fortransferring cooling fluid. The mechanism 198 may also include an innerjournal bearing extending within the arm 140 of the HP rotor shaft 34adjacent to an inside surface 208 of the arm 140 of the HP rotor shaft34, covering the plurality of circumferentially spaced and radiallyextending holes in the arm 140 of the HP rotor shaft 34. The outerjournal bearing may be fluidly connected to the outer fluid line 194 andthe inner journal bearing may be fluidly connected to the inner fluidline 196. However, in other exemplary embodiments, any other suitablemeans or mechanism 198 may be provided as the static to rotating framefluid transfer mechanism 198.

Moreover, as with other embodiments of the exemplary cooling system 100,the exemplary cooling system 100 of FIG. 8 may further include aplurality of fluid lines with at least a portion of the plurality offluid lines spaced along the circumferential direction C of the turbofanengine 10, inward of the core air flowpath 37 along the radial directionR. For example, the plurality of fluid lines may be spaced along thecircumferential direction C in substantially the same manner as theplurality of compressor fluid lines 152 are spaced along thecircumferential direction C as described above with reference to FIG. 4.Further, although the exemplary embodiment of FIG. 8 includes a singlenozzle 200 positioned adjacent to and directed towards a surface of theplurality of HP compressor rotors 106, a single nozzle 202 positionedadjacent to and directed towards a surface of the compressor dischargepressure seal 142, and a single nozzle 204 positioned adjacent to anddirected towards a surface of the HP turbine rotors 136, in otherexemplary embodiments, the exemplary cooling system 100, or rather theone or more fluid lines, may include any other suitable number ofnozzles. Alternatively, in other exemplary embodiments, the one or morefluid lines may not include one or more of the nozzles 200, 202, 204directed at the HP compressor rotors 106, the compressor dischargepressure seal 142, or the HP turbine rotors 136, respectively.

Notably, the exemplary cooling system 100 of FIG. 8 may be configured toinject cooling fluid through the inner fluid lines 196 at the same rateand/or the same ratio that the exemplary cooling system 100 describedabove with reference to FIGS. 2 through 4 is configured to providecooling fluid through the compressor fluid line(s) 152.

Moreover, it should be appreciated that in other exemplary embodiments,the second fluid line 196 of the one or more fluid lines mayadditionally, or alternatively, be configured for providing the coolingliquid indirectly to one or more of the compressor rotor 106, thecompressor discharge pressure seal 142, and/or the turbine rotor 136.For example, in other exemplary embodiments, the second fluid line 196of the one or more fluid lines may additionally, or alternatively, beconfigured for indirectly providing the cooling liquid to suchcomponents by spraying/injecting the cooling liquid into a flow ofcooling air provided over such components. Such a flow of cooling airmay be extracted from the compressor section of the turbofan engine 10.

Further, in still other exemplary embodiments of the present disclosure,the cooling system 100 may not include the inner fluid lines 196. Forexample, the one or more fluid lines of the cooling system 100 maysimply include the outer fluid line 194 fluidly connected to the staticto rotating frame fluid transfer mechanism 198, and the static torotating frame fluid transfer mechanism 198 may be configured to spraythe cooling fluid directly or indirectly on the one or more componentsof the compressor section, turbine section, or combustion section 26.With such an exemplary embodiment, the cooling system 100 may include aplurality of such mechanisms 198 positioned at any suitable locationadjacent to the one or more components for cooling.

It should additionally be appreciated that in further exemplaryembodiments of the present disclosure, aspects of the various exemplarycooling systems 100 may be combined with one another to arrive at stillother exemplary embodiments. For example, in certain exemplaryembodiments, aspects of the exemplary cooling system 100 described abovewith reference to FIGS. 2 through 4 may be combined with aspects of theexemplary cooling system 100 described above with reference to FIGS. 5and 6, and/or the exemplary cooling system 100 described above withreference to FIG. 7, and/or the exemplary cooling system 100 describedabove with reference to FIG. 8.

Inclusion of a cooling system in accordance with the exemplaryembodiments described above with reference to FIG. 7 and/or FIG. 8 in agas turbine engine may increase a lifespan of the one or more componentsof the compressor section, combustion section, or turbine section beingcooled. Additionally, or alternatively, the gas turbine engine includingsuch a cooling system may not require the components being cooled by theexemplary cooling system of FIGS. 7 and/or 8 to be formed of certainrare and/or expensive materials capable of withstanding relativelyextreme temperatures and loads. Accordingly, a cooling system inaccordance with certain exemplary embodiments of the present disclosuremay allow for more cost-efficient manufacturing of the gas turbineengine.

Referring now to FIG. 9, a method (300) for cooling a gas turbine enginein accordance with an exemplary embodiment of the present disclosure isprovided. The gas turbine engine may, in certain exemplary aspects, beconfigured as the exemplary turbofan engine 10 described above withreference FIG. 1. Accordingly, in certain exemplary aspects, the gasturbine engine may define a core air flowpath extending through acompressor section, a combustion section, and a turbine section.

The exemplary method (300) includes at (302) operating the gas turbineengine to achieve a high power output. As used herein, a “high poweroutput” refers to at least about seventy-five percent (75%) of themaximum power output of the gas turbine engine. For example, the gasturbine engine may be operated to achieve a high power output duringtakeoff and/or during climb operation modes of an aircraft including theexemplary gas turbine engine.

The exemplary method (300) also includes at (304) providing a consumablecooling liquid using an exemplary cooling system to one or more portionsof the gas turbine engine. For example, in certain exemplary aspects,providing the consumable cooling liquid using the cooling system at(304) may include providing the consumable cooling liquid to compressedair proximate an aft stage of rotor blades in the compressor sectionand/or to one or more components of the compressor section proximate theaft stage of rotor blades in the compressor section. A cooling system inaccordance with the exemplary embodiment described with reference toFIGS. 2 through 4 and/or with reference to FIGS. 5 and 6 may be used insuch an exemplary aspect.

Additionally, or alternatively, in other exemplary aspects, providingthe consumable cooling liquid using the cooling system to one or moreportions of the gas turbine engine at (304) may include providing theconsumable cooling liquid directly or indirectly to a surface of one ormore components of the compressor section, combustion section, or theturbine section not directly exposed to the core air flowpath. Forexample, in certain exemplary aspect, providing the consumable coolingliquid to one or more portions of the gas turbine engine at (304) mayinclude providing a consumable cooling liquid to one or more of an outersurface of a compressor liner, an outer surface of an outer combustionchamber liner, or to an outer surface of a turbine liner. A coolingsystem in accordance with the exemplary embodiment described above withreference to FIG. 7 may be used in such an exemplary aspect.

Additionally, or alternatively still, in certain exemplary aspects,providing the consumable cooling liquid using the cooling system to oneor more components of the compressor section, combustion section, or theturbine section at (304) may include providing the consumable liquiddirectly or indirectly to one or more of a surface of one or morecompressor rotors, a surface of one or more turbine rotors, and/or asurface of a compressor discharge pressure seal. A cooling system inaccordance with the exemplary embodiment described above with referenceto FIG. 8 may be used in such an exemplary aspect.

Notably, in at least certain exemplary aspects, providing the coolingliquid using the cooling system to one or more components of the gasturbine engine at (304) may include providing the consumable coolingliquid at the same rate and/or at the same ratio that the exemplarycooling system 100 of FIGS. 2 through 4 provides the cooling fluidthrough the exemplary compressor lines 152. For example, providing thecooling liquid using the cooling system to one or more components of thegas turbine engine at (304) may include providing up to about two poundsof consumable cooling liquid per second. Additionally, or alternatively,providing the consumable cooling liquid to one or more components of thegas turbine engine at (304) may include providing the consumable coolingliquid at a rate greater than about 0.05 percent of a mass flow rate ofthe air flowing through the core air flowpath, and less than about tenpercent (10%) of the mass flow rate of the airflow into the core airflowpath.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A gas turbine engine defining an axial direction,the gas turbine engine comprising: a compressor section; a turbinesection located downstream of the compressor section; a combustionsection positioned between the compressor section and the turbinesection, wherein the compressor section, the turbine section, and thecombustion section together define a core air flowpath; a cooling systemfor cooling one or more components not directly exposed to the core airflowpath, wherein the one or more components belongs to one or more ofthe compressor section, the turbine section, or the combustion section;the cooling system comprising a fluid tank for storing a volume ofconsumable cooling liquid; a plurality of fluid lines in fluidcommunication with the fluid tank for carrying a flow of the consumablecooling liquid and providing the consumable cooling liquid to the one ormore components not directly exposed to the core air flowpath, whereinthe plurality of lines are positioned exterior to the one or morecomponents; wherein the one or more components belongs to one or more ofthe compressor section, the turbine section, or the combustion section,and wherein the plurality of fluid lines include a supply fluid line, acompressor fluid line, and a turbine fluid line; and a valve includingan inlet fluidly coupled to the supply fluid line and configured toreceive the flow of the consumable cooling liquid from the fluid tank, afirst outlet connected to the compressor fluid line and configured todirect the flow of the consumable cooling liquid to the compressorsection, and a second outlet connected to the turbine fluid line andconfigured to direct the flow of the consumable cooling liquid to theturbine section.
 2. The gas turbine engine of claim 1, wherein thecompressor section includes a plurality of rotors, and wherein the oneor more fluid lines of the plurality of fluid lines are configured forproviding the consumable cooling liquid to a surface of the plurality ofrotors of the compressor section.
 3. The gas turbine engine of claim 1,wherein the turbine section includes a plurality of rotors, and whereinone or more fluid lines of the plurality of fluid lines are configuredfor providing the consumable cooling liquid to a surface of theplurality of rotors of the turbine section.
 4. The gas turbine engine ofclaim 1, wherein the compressor section includes a rotor, wherein theturbine section includes a rotor, wherein the rotor of the compressorsection is attached to the rotor of the turbine section through a rotorshaft, and wherein the cooling system includes a static to rotatingframe fluid transfer mechanism attached to the rotor shaft and fluidlyconnected to one or more fluid lines of the plurality of fluid lines ofthe cooling system.
 5. The gas turbine engine of claim 4, wherein one ormore fluid lines of the plurality of fluid lines are configured toprovide the consumable cooling liquid to one or both of the rotor of thecompressor section and the rotor of the turbine section through thestatic to rotating frame fluid transfer mechanism.
 6. The gas turbineengine of claim 4, wherein one or more fluid lines of the plurality offluid lines includes a first fluid line and a second fluid line, whereinthe first fluid line is fluidly connected to the second fluid linethrough the static to rotating frame fluid transfer mechanism, andwherein the second fluid line includes an outlet directed towards atleast one of the rotor of the compressor section and the rotor of theturbine section for spraying the consumable cooling liquid on the rotorof the compressor section or the rotor of the turbine section.
 7. Thegas turbine engine of claim 1, wherein the compressor section includes acompressor liner having an outer surface, and wherein an outlet of oneor more fluid lines of the plurality of fluid lines of the coolingsystem is positioned adjacent to the outer surface of the compressorliner for spraying the consumable cooling liquid on the outer surface ofthe compressor liner.
 8. The gas turbine engine of claim 1, wherein theturbine section includes a turbine liner having an outer surface, andwherein an outlet of one or more fluid lines of the plurality of fluidlines of the cooling system is positioned adjacent to the outer surfaceof the turbine liner for spraying the consumable cooling liquid on theouter surface of the turbine liner.
 9. The gas turbine engine of claim1, wherein the combustion section includes an outer combustion chamberliner having an outer surface, and wherein an outlet of one or morefluid lines of the plurality of fluid lines of the cooling system ispositioned adjacent to the outer surface of the outer combustion chamberliner for spraying the consumable cooling liquid on the outer surface ofthe outer combustion chamber liner.
 10. The gas turbine engine of claim1, wherein the cooling system further includes a fluid pump forgenerating the flow of the consumable cooling fluid from the fluid tankthrough the plurality fluid lines.
 11. The gas turbine engine of claim1, wherein at least a portion of each of the plurality of fluid lines iscircumferentially spaced within the gas turbine engine.
 12. The gasturbine engine of claim 1, wherein the cooling system is configured toinject up to two (2) pounds of the consumable cooling liquid per secondinto a compressed air proximate to an aft-stage of rotor blades.
 13. Thegas turbine engine of claim 1, wherein the cooling system is configuredto inject the consumable cooling liquid at a rate greater than about0.05% of a mass flow rate of the air flowing through the core airflowpath and less than about ten percent of the mass flow rate of theair flowing through the core air flowpath.
 14. The gas turbine engine ofclaim 1, wherein the gas turbine engine further defines a radialdirection, and wherein the fluid tank is positioned outward of the coreair flowpath along the radial direction.
 15. The gas turbine engine ofclaim 1, wherein one or more fluid lines of the plurality of fluid linesare configured for providing the consumable cooling liquid to the one ormore components of the compressor section, the turbine section, and thecombustion section not directly exposed to the core air flowpath byproviding the consumable cooling liquid to cooling air flowing over theone or more components of the compressor section, the turbine section,and the combustion section not directly exposed to the core airflowpath.
 16. The gas turbine engine of claim 1, wherein the coolingsystem operates to cool the compressor section during periods of peakpower of the gas turbine engine.
 17. A method for cooling a gas turbineengine according to claim 1, the method comprising: operating the gasturbine engine to achieve a high power output; and using the coolingsystem to provide the consumable cooling liquid to a surface of the oneor more components of the compressor section, the combustion section, orthe turbine section, the surface not directly exposed to the core airflowpath.
 18. The method of claim 17, wherein the cooling liquid is atleast one of water and a glycol-water mixture.
 19. The method of claim17, wherein providing the consumable cooling liquid using the coolingsystem to the surface of the one or more components in the compressorsection, the combustion section, or the turbine section includesproviding up to two (2) pounds of the consumable cooling liquid persecond into a compressed air proximate to an aft-stage of rotor blades.20. The method of claim 17, wherein providing the consumable coolingliquid using the cooling system to the surface of the one or morecomponents in the compressor section, the combustion section, or theturbine section includes providing the consumable cooling liquid at arate greater than about 0.05% of a mass flow rate of air flowing throughthe core air flowpath and less than about ten percent of the mass flowrate of the air flowing through the core air flowpath.